A <span class="c5 g0">turbinespan> <span class="c6 g0">rotorspan> <span class="c7 g0">bladespan> with a row of exit slots adjacent to a <span class="c10 g0">trailingspan> <span class="c11 g0">edgespan> region of the airfoil that extends from the platform to the <span class="c7 g0">bladespan> tip, where the row of exit slots is formed into three groups that include an upper span group, a mid-span group and a lower span group. The upper span group discharges <span class="c15 g0">filmspan> <span class="c16 g0">coolingspan> air upward with respect to the airfoil <span class="c0 g0">chordwisespan> <span class="c1 g0">directionspan>, the mid-span group discharges along the <span class="c0 g0">chordwisespan> <span class="c1 g0">directionspan>, and the lower span group discharges in a <span class="c1 g0">directionspan> downward with respect to the <span class="c0 g0">chordwisespan> <span class="c1 g0">directionspan>. Each exit slot is formed with a plurality of metering inlet holes that discharge into a first diffusion section and then a second diffusion section.

Patent
   8231330
Priority
May 15 2009
Filed
May 15 2009
Issued
Jul 31 2012
Expiry
Sep 06 2030
Extension
479 days
Assg.orig
Entity
Small
12
13
EXPIRED
1. An air cooled <span class="c5 g0">turbinespan> <span class="c6 g0">rotorspan> <span class="c7 g0">bladespan> comprising:
an airfoil having a leading <span class="c11 g0">edgespan> and a <span class="c10 g0">trailingspan> <span class="c11 g0">edgespan>;
the airfoil having a pressure side wall and a suction side wall both extending between the leading <span class="c11 g0">edgespan> and the <span class="c10 g0">trailingspan> <span class="c11 g0">edgespan>;
a row of exit slots arranged along the pressure side wall of the airfoil adjacent to the <span class="c10 g0">trailingspan> <span class="c11 g0">edgespan> region of the airfoil;
the row of exit slots extending from adjacent to a platform of the <span class="c7 g0">bladespan> to adjacent to a tip of the <span class="c7 g0">bladespan>;
the row of exit slots including an upper span section of slots, a mid-span section of slots and a lower span section of slot; and,
the upper span exit slots discharging <span class="c15 g0">filmspan> <span class="c16 g0">coolingspan> air in a <span class="c1 g0">directionspan> toward the <span class="c7 g0">bladespan> tip, the mid-span exit slots discharging <span class="c15 g0">filmspan> <span class="c16 g0">coolingspan> air in a <span class="c0 g0">chordwisespan> <span class="c1 g0">directionspan> of the airfoil, and the lower span exit slots discharging <span class="c15 g0">filmspan> <span class="c16 g0">coolingspan> air in a <span class="c1 g0">directionspan> toward the platform.
2. The air cooled <span class="c5 g0">turbinespan> <span class="c6 g0">rotorspan> <span class="c7 g0">bladespan> of claim 1, and further comprising:
the exit slots each include a metering inlet section with a plurality of metering holes that open into a diffusion section.
3. The air cooled <span class="c5 g0">turbinespan> <span class="c6 g0">rotorspan> <span class="c7 g0">bladespan> of claim 2, and further comprising:
the diffusion section includes a first diffusion section immediately downstream from the metering section and a second diffusion section immediately downstream from the first diffusion section.
4. The air cooled <span class="c5 g0">turbinespan> <span class="c6 g0">rotorspan> <span class="c7 g0">bladespan> of claim 3, and further comprising:
the first diffusion section includes an expansion of from 7 to 13 degrees on a downstream wall with respect to an axis of the metering holes.
5. The air cooled <span class="c5 g0">turbinespan> <span class="c6 g0">rotorspan> <span class="c7 g0">bladespan> of claim 4, and further comprising:
the upper span exit slots have a radial outboard expansion of 0 to 3 degrees and a radial inboard expansion of from 7 to 13 degrees; and,
the mid-span exit slots have a radial outboard expansion of 7 to 13 and a radial inboard expansion of from 7 to 13 degrees; and,
the lower-span exit slots have a radial outboard expansion of 7 to 13 and a radial inboard expansion of from 0 to 3 degrees.
6. The air cooled <span class="c5 g0">turbinespan> <span class="c6 g0">rotorspan> <span class="c7 g0">bladespan> of claim 1, and further comprising:
the upper span exit slots have a radial outboard expansion of 0 to 3 degrees and a radial inboard expansion of from 7 to 13 degrees; and,
the mid-span exit slots have a radial outboard expansion of 7 to 13 and a radial inboard expansion of from 7 to 13 degrees; and,
the lower-span exit slots have a radial outboard expansion of 7 to 13 and a radial inboard expansion of from 0 to 3 degrees.
7. The air cooled <span class="c5 g0">turbinespan> <span class="c6 g0">rotorspan> <span class="c7 g0">bladespan> of claim 1, and further comprising:
the upper span exit slots have a discharge angle with respect to the airfoil spanwise radial <span class="c1 g0">directionspan> of from 20 to 40 degrees;
the mid-chord span exit slots have a discharge angle with respect to the airfoil spanwise radial <span class="c1 g0">directionspan> of from 80 to 100 degrees; and,
the lower span exit slots have a discharge angle with respect to the airfoil spanwise radial <span class="c1 g0">directionspan> of from 110 to 130 degrees.
8. The air cooled <span class="c5 g0">turbinespan> <span class="c6 g0">rotorspan> <span class="c7 g0">bladespan> of claim 1, and further comprising:
the upper span exit slots have a discharge angle with respect to the airfoil spanwise radial <span class="c1 g0">directionspan> of around 30 degrees;
the mid-chord span exit slots have a discharge angle with respect to the airfoil spanwise radial <span class="c1 g0">directionspan> of around 90 degrees; and,
the lower span exit slots have a discharge angle with respect to the airfoil spanwise radial <span class="c1 g0">directionspan> of around 120 degrees.

None.

None.

1. Field of the Invention

The present invention relates generally to a gas turbine engine, and more specifically to a turbine rotor blade with film cooling slots.

2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98

A gas turbine engine, such as an industrial gas turbine (IGT) engine, includes a turbine with multiple rows or stages or stator vanes that guide a high temperature gas flow through adjacent rotors of rotor blades to produce mechanical power and drive a bypass fan, in the case of an aero engine, or an electric generator, in the case of an IGT. In both cases, the turbine is also used to drive the compressor.

It is well known in the art of gas turbine engine design that the efficiency of the engine can be increased by passing a higher gas flow temperature through the turbine. However, the turbine inlet temperature is limited by the material properties of the turbine, especially for the first stage airfoils since these are exposed to the highest temperature gas flow. As the gas flow passes through the various stages of the turbine, the temperature decreases as the energy is extracted by the rotor blades.

Another method of increases the turbine inlet temperature is to provide more effective cooling of the airfoils. Complex internal and external cooling circuits or designs have been proposed using a combination of internal convection and impingement cooling along with external film cooling to transfer heat away from the metal and form a layer of protective air to limit thermal heat transfer to the metal airfoil surface. However, since the pressurized air used for the airfoil cooling is bled off from the compressor, this bleed off air decreases the efficiency of the engine because the work required to compress the air is not used for power production. It is therefore wasted energy as far as producing useful work in the turbine.

Shaped diffusion film cooling holes are normally used for the cooling of a turbine blade pressure side wall. The use of axial oriented film cooling holes on the pressure side surface of the airfoil is mainly for an injection of cooling air inline with the main stream gas flow which is accelerated into multiple directions as represented by the various arrows in FIG. 1.

However, at the airfoil pressure side surface two thirds of the way downstream from the leading edge region, the hot gas secondary flow migrates in the multiple directions, depending on the pressure gradient and also moving in an axial direction. Due to pressure gradient across the blade tip, the upper blade span hot gas flow migrates toward the blade tip section. Due to the nature of turbine expansion, the middle portion flow in the axial direction. Due to the hot gas passage channel pressure gradient, the lower span hot gas flow migrates toward the blade platform. An axial oriented shaped film cooling hole is used in that region of the blade thus becomes undesirable. FIG. 1 shows the secondary hot gas flow phenomena on the blade pressure side surface.

It is an object of the present invention to provide for a turbine rotor blade with a film cooling slot arrangement that will reduce the migration of hot gas flow toward the blade tip over that of the cited prior art blades.

It is another object of the present invention to provide for a turbine rotor blade with an improved blade tip cooling design so that the blade life is increased over that of the cited prior art blades.

These objectives and more can be achieved by the turbine rotor blade of the present invention that includes a row of film cooling slots located on the pressure side wall and adjacent to a trailing edge region of the airfoil in which the slots are multiple compound angled multi-diffusion film cooling slots at a special spanwise angle relative to the airfoil. The slots are formed into three equal groups in the row and include an lower span group with a discharge angle oriented downward from a hot gas flow direction, a mid-span group with a discharge angle oriented parallel to the hot gas flow path, and an upper spanwise group with a discharge angle oriented upward from the hot gas flow path.

Each slot includes multiple metering holes that open into a first diffusion chamber and then into a second diffusion chamber before discharging out from the slot opening. With this design, the compound angled multi-diffusion film cooling slots allow the cooling air flow to discharge from each individual metering hole to be injected onto the airfoil surface at a certain spanwise angle and to be diffused within the diffuser. This yields a good buildup of the coolant sub-boundary layer next to the airfoil pressure side surface to form a film layer without shear mixing effect to seal the airfoil fro the hot gas flow.

FIG. 1 shows a prior art turbine blade with the secondary hot gas flow migration on the pressure side wall toward the blade tip.

FIG. 2 shows a cross section view of one of the compound angled multi-diffusion film cooling slots of the present invention.

FIG. 3 shows a top view of the compound angled multi-diffusion film cooling slots of the present invention in FIG. 2.

FIG. 4 shows a turbine rotor blade with a row of the compound angled multi-diffusion film cooling slots of the present invention.

The present invention is a turbine rotor blade with a row of film cooling slots located adjacent to the trailing edge region of the blade to reduce or eliminate the prior art hot gas flow migration problem described above in the prior art toward the blade tip. The blade of the present invention is intended for use in an industrial gas turbine engine, but can be adapted for use in an aero engine. FIG. 2 shows a cross section side view of one of the compound angled multi-diffusion film cooling slots 10 used on the blade of the present invention. The slot 10 includes a metering inlet section 11 of a constant diameter, a first expansion section located immediately downstream from the metering section 11, and a second expansion section 13 located immediately downstream from the first expansion section 12. The first expansion section includes a downstream wall with an expansion of from 7 to 13 degrees with respect to the axis of the metering hole in the metering section 11. The second expansion section 13 includes a downstream wall with an additional expansion of 7 to 13 degrees with respect to the downstream wall in the first expansion section 12. The second diffusion section 13 opens onto the outer airfoil surface of the airfoil wall 9.

FIG. 3 shows a top view of the slot 10 of FIG. 2 with a number of metering holes 21 each of a constant diameter that forms the metering section 11 of the slot 10. In the various embodiments, the number of metering holes for each slot can be from 3 to 5 or 6. The first expansion section 12 also includes two side walls that have an expansion each of the side walls but vary in the expansion angle depepending upon the location of the slot within the three zones or groups. The second expansion section also includes two side walls with an expansion that varies depending upon which group the slot is in.

FIG. 4 shows the blade with a row of 6 slots being divided up into a lower span group nearest to the platform, a middle span group and an upper span group nearest to the blade tip. Each of the three groups takes up around one third of the spanwise distance of the airfoil so that they form equal spanwise length groups to cover the three spanwise sections of the airfoil.

Each of the three slot groups is oriented at a different spanwise angle relative to the blade. The lower span group is at 110 to 130 degrees spanwise angle and located in the lower one third of the spanwise length of the airfoil. The mid-span group is oriented at 80 to 100 degrees spanwise angle and located in the middle one third of the spanwise length of the airfoil. The upper span group is at 20 to 40 degrees from the spanwise angle and is located in the upper one third of the spanwise length of the airfoil. All of the slots can be formed by EDM or laser drilling the two diffusion sections onto the airfoil pressure side wall followed by drilling the multi-metering holes into each individual diffusion slot. The spanwise angle is the angle defined between the blade radial span direction to the centerline of the film cooling hole in a clockwise rotation direction. A line parallel to the airfoil spanwise (radial) direction will be at zero degrees while the direction parallel to the chordwise direction of the airfoil (and in an aft direction) will be at 90 degrees.

The main purpose of the compound angled multi-diffusion film cooling slots is to allow the cooling flow discharged from each individual metering hole to be ejected onto the airfoil surface at a specific spanwise angle and diffused within the diffuser. This yields a good buildup of the coolant sub-boundary layer next to the airfoil pressure side surface and forms a layer of film cooling air without the shear mixing effect in order to better seal the airfoil wall from the hot gas flow.

Each of the slots is formed of two main portions: a first portion for the metering holes which are at a constant diameter. These metering holes are drilled at the same orientation as the compound angled multi-diffusion film cooling slot. The second portion is the multi-diffusion slot which is shaped depending upon which spanwise group it is in. the upper span group of slots 10 is formed with a 0-3 degree expansion in the spanwise radial outward direction (top wall surface as seen in FIG. 4). The multiple expansion concept is incorporated in the spanwise radial direction, a 7-13 degree first expansion from the end of the metering hole to the diffuser exit plane followed by a second expansion of 7-13 degree from the diffuser exit plane to the airfoil exterior surface. All of these expansion angles are relative to the centerline of the metering hole.

In the mid-span group, the multi-diffusion slot is formed with a 7-13 degree expansion in the spanwise radial outward and inboard directions for the first expansion. Thus, both the top wall and the bottom wall as seen in FIG. 4 for the slot expands from 7-13 degrees. The second expansion is 7-13 degrees from the diffuser exit plane to the airfoil exterior surface. All of these expansion angles are relative to the centerline of the metering hole. All of the diffusion angles are relative to the centerline of the metering hole.

The lower span group includes a 0-3 degrees expansion in the spanwise radial inboard direction (the bottom wall as seen in FIG. 4). The multiple expansion concept is incorporated into the spanwise radial outward direction with a 7-13 degree first expansion from the end of the metering hole to the diffuser exit plane flowed by a second expansion of 7-13 degrees from the diffuser exit plane to the airfoil exterior surface. The lower span slots are a mirror image of the upper span slots.

Thus, the upper span slots discharge in a direction in the range of 20-40 degrees from the radial spanwise direction of the airfoil, preferably at around 30 degrees; the mid-span slots discharge in the range of 80-100 degrees from the radial spanwise direction of the airfoil, preferably at around 90 degrees; and the lower span slots discharge in a direction in the range of 110-130 degrees from the radial spanwise direction of the airfoil, preferably at around 120 degrees.

Liang, George

Patent Priority Assignee Title
10030525, Mar 18 2015 General Electric Company Turbine engine component with diffuser holes
10174620, Oct 15 2015 General Electric Company Turbine blade
10400609, Jun 21 2012 RTX CORPORATION Airfoil cooling circuits
10641103, Jan 19 2017 RTX CORPORATION Trailing edge configuration with cast slots and drilled filmholes
10808551, Jun 21 2012 RTX CORPORATION Airfoil cooling circuits
10815788, Jan 30 2017 RTX CORPORATION Turbine blade with slot film cooling
11021969, Oct 15 2015 General Electric Company Turbine blade
11401821, Oct 15 2015 General Electric Company Turbine blade
9316104, Oct 25 2012 RAYTHEON TECHNOLOGIES CORPORATION Film cooling channel array having anti-vortex properties
9360020, Apr 23 2014 Electric Torque Machines Inc Self-cooling fan assembly
9752440, May 29 2015 General Electric Company Turbine component having surface cooling channels and method of forming same
9879546, Jun 21 2012 RTX CORPORATION Airfoil cooling circuits
Patent Priority Assignee Title
3781129,
4940388, Dec 07 1988 Rolls-Royce plc Cooling of turbine blades
5096379, Oct 12 1988 Rolls-Royce plc Film cooled components
5356265, Aug 25 1992 General Electric Company Chordally bifurcated turbine blade
5846057, Dec 12 1995 General Electric Company Laser shock peening for gas turbine engine weld repair
6494678, May 31 2001 General Electric Company Film cooled blade tip
6514037, Sep 26 2001 General Electric Company Method for reducing cooled turbine element stress and element made thereby
7021896, May 23 2003 Rolls-Royce plc Turbine blade
7249934, Aug 31 2005 General Electric Company Pattern cooled turbine airfoil
20020182074,
20070286729,
20090081048,
20090180861,
/////////////
Executed onAssignorAssigneeConveyanceFrameReelDoc
May 15 2009Florida Turbine Technologies, Inc.(assignment on the face of the patent)
Jul 18 2012LIANG, GEORGEFLORIDA TURBINE TECHNOLOGIES, INCASSIGNMENT OF ASSIGNORS INTEREST SEE DOCUMENT FOR DETAILS 0287310926 pdf
Mar 01 2019FLORIDA TURBINE TECHNOLOGIES INC SUNTRUST BANKSUPPLEMENT NO 1 TO AMENDED AND RESTATED INTELLECTUAL PROPERTY SECURITY AGREEMENT0485210081 pdf
Mar 01 2019S&J DESIGN LLCSUNTRUST BANKSUPPLEMENT NO 1 TO AMENDED AND RESTATED INTELLECTUAL PROPERTY SECURITY AGREEMENT0485210081 pdf
Mar 01 2019CONSOLIDATED TURBINE SPECIALISTS LLCSUNTRUST BANKSUPPLEMENT NO 1 TO AMENDED AND RESTATED INTELLECTUAL PROPERTY SECURITY AGREEMENT0485210081 pdf
Mar 01 2019ELWOOD INVESTMENTS LLCSUNTRUST BANKSUPPLEMENT NO 1 TO AMENDED AND RESTATED INTELLECTUAL PROPERTY SECURITY AGREEMENT0485210081 pdf
Mar 01 2019TURBINE EXPORT, INC SUNTRUST BANKSUPPLEMENT NO 1 TO AMENDED AND RESTATED INTELLECTUAL PROPERTY SECURITY AGREEMENT0485210081 pdf
Mar 01 2019FTT AMERICA, LLCSUNTRUST BANKSUPPLEMENT NO 1 TO AMENDED AND RESTATED INTELLECTUAL PROPERTY SECURITY AGREEMENT0485210081 pdf
Mar 01 2019KTT CORE, INC SUNTRUST BANKSUPPLEMENT NO 1 TO AMENDED AND RESTATED INTELLECTUAL PROPERTY SECURITY AGREEMENT0485210081 pdf
Mar 30 2022TRUIST BANK AS SUCCESSOR BY MERGER TO SUNTRUST BANK , COLLATERAL AGENTKTT CORE, INC RELEASE BY SECURED PARTY SEE DOCUMENT FOR DETAILS 0596190336 pdf
Mar 30 2022TRUIST BANK AS SUCCESSOR BY MERGER TO SUNTRUST BANK , COLLATERAL AGENTFTT AMERICA, LLCRELEASE BY SECURED PARTY SEE DOCUMENT FOR DETAILS 0596190336 pdf
Mar 30 2022TRUIST BANK AS SUCCESSOR BY MERGER TO SUNTRUST BANK , COLLATERAL AGENTCONSOLIDATED TURBINE SPECIALISTS, LLCRELEASE BY SECURED PARTY SEE DOCUMENT FOR DETAILS 0596190336 pdf
Mar 30 2022TRUIST BANK AS SUCCESSOR BY MERGER TO SUNTRUST BANK , COLLATERAL AGENTFLORIDA TURBINE TECHNOLOGIES, INCRELEASE BY SECURED PARTY SEE DOCUMENT FOR DETAILS 0596190336 pdf
Date Maintenance Fee Events
Dec 10 2015M2551: Payment of Maintenance Fee, 4th Yr, Small Entity.
Mar 23 2020REM: Maintenance Fee Reminder Mailed.
Sep 07 2020EXP: Patent Expired for Failure to Pay Maintenance Fees.


Date Maintenance Schedule
Jul 31 20154 years fee payment window open
Jan 31 20166 months grace period start (w surcharge)
Jul 31 2016patent expiry (for year 4)
Jul 31 20182 years to revive unintentionally abandoned end. (for year 4)
Jul 31 20198 years fee payment window open
Jan 31 20206 months grace period start (w surcharge)
Jul 31 2020patent expiry (for year 8)
Jul 31 20222 years to revive unintentionally abandoned end. (for year 8)
Jul 31 202312 years fee payment window open
Jan 31 20246 months grace period start (w surcharge)
Jul 31 2024patent expiry (for year 12)
Jul 31 20262 years to revive unintentionally abandoned end. (for year 12)